Gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine vane and blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for additional thermal protection.
Typically, turbine vanes are formed from an elongated portion forming an airfoil having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to an inner endwall. The turbine vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. Additionally, the turbine vane includes an outer diameter endwall at a first end and an inner diameter endwall at a second end. The inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system. These cooling circuits in the vanes receive air from the compressor of the turbine engine and pass the air through the ends of the vane adapted to be coupled to the vane carrier. The cooling circuits often include multiple flow paths that are designed to maintain all areas of the turbine vane at a relatively uniform temperature. At least some of the air passing through these cooling circuits may be exhausted through orifices in the wall of the vane.
U.S. Pat. No. 6,955,523 to McClelland discloses such a cooling circuit including a serpentine network of channels passing between the suction and pressure sides of the turbine vane, where each channel extends between turns of the serpentine network positioned at the inner diameter and outer diameter endwalls.
U.S. Patent Application Publication No. 2005/0244270 to the inventor of the present invention, discloses a cooling circuit for a turbine blade including channels within the suction and pressure sides for passing cooling fluid toward the turbine blade tip at the first end for creating a counterflow to a leakage flow of combustor gases between the blade tip and an outer seal.
An additional cooling system for a turbine blade is disclosed in U.S. Patent Application Publication No. 2005/0031452, also to the inventor of the present invention, and discloses directing cooling fluid into a center cavity between the pressure and suction sides, after which the cooling fluid flows through supply orifices and into cavities within the suction and pressure walls for spiral fluid flow before exiting the turbine blade through exhaust orifices in the outer surface of the pressure and suction sides.